Stall, buffeting, low speed and high attitude protection system

ABSTRACT

A flight control system moves elevators according to a pilot command summed with an automatic command. The flight control system monitors a set of flight parameters to determine if the flight vehicle is operating inside a permitted envelope. The flight controls system incorporates automatic protections thru the automatic elevator command if the flight vehicle is close to its envelope limits. The exemplary illustrative non-limiting implementation herein provides automatic protections in order to protect the flight vehicle from low speeds, high attitude, stalls and buffetings.

TECHNOLOGICAL FIELD

The technology herein relates to a flight control system for anaircraft. More particularly, the technology herein relates to methodsand apparatus for providing multiple protections to an aircraft equippedwith an inceptor for inputting pilot commands.

BACKGROUND AND SUMMARY

While man has mastered travel on land, sea and air, there still exists,in travel through or across any of the three, some risk of peril.Although trained drivers, captains and pilots may have worked for yearsto develop their skill, human error still happens. Further, in thepresence of adverse conditions such as bad weather, slight errors ormiscalculations may be exacerbated into highly dangerous ones.

As technology advances, computers play a much more active role in aidingvehicle maneuvering. Features such as traction control, braking andsteering are often processed at least in part by a car computer chip,for example, and more sophisticated car computers can even detectadverse weather conditions and compensate to help keep a driver safe.

Similarly, the use of feedback control laws to augment the pitch commandof an aircraft has been used since the latter half of 20^(th) Century.In terms of modern aircraft, digital control laws are used to implementcontrol laws that use a reference command based on pitch rate, loadfactor or a combination of thereof. Airspeed in conjunction with a loadfactor may also be considered as a reference command. In some cases, allthree variables are considered as reference command, that is, the loadfactor, pitch rate and airspeed are considered.

The exemplary illustrative non-limiting implementations provide furthersafety controls for aircraft. For example, the flight control law of oneexemplary illustrative non-limiting implementation computes anaugmentation command correction based on a set of flight parameters andon the sensed position of the pilot inceptor. The pilot inceptor may beany of a plurality of devices used in aeronautics industry to serve asan interface with a human pilot, e.g. columns, mini-columns, sticks,control yokes, side-sticks, etc. The augmentation command may be mixedwith a direct mode pilot command, which may be sent straight to thepitch control surface actuator. The actuator controls a pitch controlsurface such as an elevator.

Just as driver operations may be altered by a computer chip in a car toprevent accidents on the road, the augmentation command may performstability augmentation with some additional protection functions for anairplane, which are designed to avoid some undesirable events, such as:i) stall, ii) stall with icing, iii) buffeting, iv) horizontalstabilizer high load, v) low speed, vi) high pitch attitude, etc.

According to one exemplary illustrative non-limiting implementation, thecontrol law computes a reference command (δ_(law)), in degrees, which isbased at least in part on the position of the pilot inceptor. Thisfunction is called command shaping, and the function may change duringflight. This reference command may be used both in feed-forward andintegral loops: the feed-forward command may be produced based at leastin part on a gain multiplied by the reference command (δ_(law)); theintegral command may be based at least in part on the integral of theerror difference between either angle of attack (α), or pitch angle (θ),and the reference command (δ_(law)), multiplied by another gain. Thus,for example, the error may be e=δ_(law)−α or e=δ_(law)−θ.

Further, in this exemplary illustrative non-limiting implementation, thefeedback loop may also consider a state feedback based on a set ofsensed flight parameters such as angle of attack (α), pitch rate (q),pitch angle (θ) and airspeed (u) which may be combined using a set ofgains.

The integral, feedback and feed-forward command may be summed tocompound the augmentation command, which drives the pitch control andtrends to reduce the error e to zero in steady state due to integralfeedback.

According to this exemplary implementation, the gains may be computedsuch that the command augmentation automatically pitches the airplanedown when one or more undesirable conditions, such as the ones mentionedabove, are detected.

According to a further exemplary illustrative non-limitingimplementation, based at least in part on a set of flight parameters, alogic module may be at least partly responsible to define the engagementof a control law in a protection function, such as those mentioned,which may be made dynamically during the flight. In a given flightcondition, depending on the protection function performed, the logicmodule may change the following in the control law: i) all the gains ofthe control law, ii) the command shaping function that defines therelation between pilot command and reference command δ_(law), and/oriii) switch selection between angle of attack (α) or pitch angle (θ) inthe integral command. In this exemplary implementation, the shapingfunction defines a maximum commanded angle of attack or pitch angle,correspondent to a maximum inceptor position, depending on which of themare being fed back in a given instant. This way, it is possible to limitthe aircraft envelope as desired, using the same law structure thatserves as a variety of protections, in different flight phases.

To define all that, the logic module and command shaping uses a set ofparameters, which comprises: height above ground (h_(AGL)), icedetection bit (b_(Ice)) and engine throttle lever position (δ_(TLA)).

When the logic module is not engaged, this control law may not send anycommand; i.e., a null augmentation command may be sent.

Also, the gains may change depending on flight envelope parameters andconfiguration parameters, such as Mach number, altitude, flap positionand landing gear position.

Thus, according to one exemplary illustrative non-limitingimplementation, either angle of attack or pitch angle are considered asa reference command. Further, the angle of attack and/or pitch anglevalues are limited inside a permitted flight envelope by means of acommand shaping, and gains are changed, adapting one or more protectionfunctions.

In accordance with another exemplary illustrative non-limitingimplementation the command shaping, the feedback and feed-forward gainsand switches and the integral feedback from angle-of-attack (α) to pitchangle (θ) are changed.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features and advantages will be better and morecompletely understood by referring to the following detailed descriptionof exemplary non-limiting illustrative implementations in conjunctionwith the drawings of which:

FIG. 1 is an example of one flight vehicle—a civil transporterturbo-fan;

FIG. 2 is a schematic diagram of an exemplary illustrative non-limitingflight control system, showing the basic architecture of the system;

FIG. 3 presents a schematic of the exemplary illustrative non-limitingsoftware that processes the function of the exemplary flight controlsystem, showing how the pilot command is transformed into an elevatorcommand depending on a logic module to enable it;

FIG. 4 is a diagram that details the exemplary illustrative non-limitinglogic module, which enables the elevator command based on a series ofsensor inputs; and

FIG. 5 is an exemplary logic flow of an exemplary routine according toone non-limiting implementation.

DETAILED DESCRIPTION

The exemplary illustrative non-limiting implementations herein relate tosystems, apparatuses and methods to be used in a flight vehicle equippedwith pitch control, such as elevators and a pilot inceptor such as aside-stick or a column yoke. FIG. 1 shows an exemplary illustrative atwin turbo-fan engine 114 civilian transporter aircraft. The plane has aset of wings 113, provided with spoilers 112 and flaps 116. The spoilers112 help change lift, drag and roll, and the flaps 116 help change liftand drag. The tail of the plane is also equipped with a horizontalstabilizer 117 provided with an elevator 115 which controls pitchorientation of the aircraft in flight.

An exemplary illustrative non-limiting flight control system is shown inFIG. 2. This exemplary flight control system receives input positionsignals from the pilot inceptor 202 command (p). The term “pilotinceptor” includes a plurality of devices used in aeronautics industryto allow the interface with the human pilot, e.g. columns, mini-columns,sticks, side-sticks and all others.

Further, the exemplary illustrative non-limiting system receives signalsfrom a set of sensors 218, 219, 220, 221, 222. In this exemplaryimplementation, the sensors provide: angle-of-attack (α),angle-of-attack rate ({dot over (α)}), airspeed (u), airspeed rate ({dotover (u)}), the flap position (δ_(F)), gear position (δ_(G)), pitchattitude (θ), pitch rate (q), height above ground (h_(AGL)), icedetection bit (b_(Ice)), engine throttle lever position (δ_(TLA)), Machnumber (Mach) and altitude (h). Other sensors are also possible.

According to this exemplary implementation, the information flows via ameans of transmitting multiple data such as a bus 205. All the data,i.e. pilot commands and sensors, is sent to a processor 204 that isoperable to compute output based, for example, on a programmable code.The processor 204 is able, for example, to compute an elevator commandbased on the input data received. This command is sent to a mechanism toactuate a flight control surface 207, which comprises a control unitable to command the elevator surfaces 201 to the commanded position.Resultantly, the elevator surfaces are deployed according to the commandcomputed by the processor 204.

FIG. 3 shows exemplary main units of exemplary illustrative non-limitingsoftware that may process the function in an exemplary flight controlssystem. The pilot command block 305 represents the position of the pilotinceptor which is sent directly to the elevator surface 306. Accordingto this exemplary implementation, as long as the stall, buffeting, highattitude or low speed protections are active, this inceptor command iscancelled, i.e. the aircraft is completely controlled in the pitch axisthrough the full authority automatic system.

In the exemplary illustrative non-limiting implementation, pilotinceptor command is transformed into alpha (α) command when the stall,buffeting and low speed protections are active or into pitch angle (θ)command when high attitude protection is activated. The relation betweenthe variable to be controlled (α or θ) and pilot command is depicted ascommand shaping 308. The output of the command shaping (δ_(law)) is usedas reference to manipulate the elevators to track the variables α or θ.When the pilot moves the inceptor to the stop (i.e. the mechanical limitof the inceptor), command shaping produces a maximum α or θ in order topreclude the airplane from exceeding the maximum allowed α or θ for thecurrent airplane configuration.

The state feedback, feed-forward command and integral command compoundthe automatic elevator command. The state feedback signal is calculatedusing the pitch states of the aircraft dynamic 307 which are fed back tothe closed loop control law. Airspeed (u), pitch rate (q), pitch angle(θ) and angle of attack (α) are multiplied by the gains listed as 301,302, 303, 304, respectively. The feed-forward command is produced basedon the feed-forward gain 309 multiplied by the reference generated bythe command shaping output 308.

The error (e) is calculated as the result of the difference between thereference and the angle-of-attack or pitch angle. The angle-of-attack isused when the stall, low speed and/or buffeting protections are engaged.The pitch angle (θ) is used when the high attitude protection isengaged. The integral of the error (e) is multiplied by the integralgain in order to produce the integral command.

The gains values depend upon which protection is active. For example,when the low speed protection is active, the pitch angle gain 303 andtrue airspeed gain 301 are increased when compared to the pitch anglegain 303 and true airspeed gain 301 used in the stall protectionfunction. Also, the gains are scheduled according to the Mach number andaltitude the airplane is flying at the moment the protection is engaged.

FIG. 4 comprises all data processing according to one exemplaryillustrative implementation to allow the proper engagement and gainswitching of the exemplary flight control system mode, according toflight conditions.

The automatic elevator command may enabled when any of the conditionsbelow is true:

-   -   1. The angle-of-attack plus a bias based at least in part on the        angle-of-attack rate is above the angle-of-attack reference        value.    -   2. The airspeed minus a bias based at least in part on the        airspeed rate is below the airspeed reference value.    -   3. The pitch attitude plus a bias based at least in part on the        pitch rate is above the pitch attitude reference value.

The angle-of-attack reference value depends at least in part upon theMach number, landing gear, flap position and ice condition. The airspeedreference value depends at least in part upon the flap position. Thepitch attitude reference value depends at least in part upon the flapposition and height above ground level. The height above ground level isestimated based at least in part on ground speed and flight path anglefor take-off and based on radar altimeter sensors for landing.

FIG. 5 shows an exemplary flow of an algorithm for determiningprocessing of an inceptor command. Initially, the inceptor data isobtained (step 501). Along with the inceptor data, sensor data isobtained (step 503). The sensor data can be used to determine a varietyof flight parameters, and can further be used to determine if any of theprotections are active. Based on the sensor data, a system running theexemplary algorithm will determine if stall, buffeting or low speedprotections are active (step 505). If any of those are active, then theinceptor command, in this exemplary implementation, is transformed intoangle-of-attack reference a (step 511). If none of the mentionedprotections are active, then the system determines if high attitudeprotection is active (step 507). If high attitude protection is active,then the inceptor command is transformed into pitch angle reference θ(step 513). If no protections are active, then the inceptor command isprocessed directly (step 509).

In the cases where protections are active, the system provides commandshaping to transform the initial inceptor command into a control commandfor the elevator. To do so, the system calculates state feedback,feed-forward and/or integral commands and applies them through atransformation function to the respective reference values α or θ (steps515, 517). Then, an elevator command is output to an actuator (step 519)and the actuator adjusts the elevator (step 521).

While the technology herein has been described in connection withexemplary illustrative non-limiting implementations, the invention isnot to be limited by the disclosure. The invention is intended to bedefined by the claims and to cover all corresponding and equivalentarrangements whether or not specifically disclosed herein.

1. A flight control apparatus operable to deploy the control surfaces ofa flight vehicle to create automatic protections in order to protect theflight vehicle, comprising: a processor; an inceptor including a commandsensor to sense and deliver inceptor commands to the processor; aplurality of sensors to sense and deliver a plurality of parameters tothe processor; a first transformation routine associated with theprocessor to transform the inceptor commands; a second transformationroutine associated with the processor to transform the transformedinceptor commands into output commands; a determination routineassociated with the processor to determine whether a protection isengaged; and an actuator driven by the output commands to control atleast one control surface, wherein said determination routine is capableof determining whether a stall protection, buffeting protection, lowspeed protection and/or high attitude protection is engaged; wherein ifa stall protection, buffeting protection or low speed protection isengaged, said first transformation routine is operable to transform saidinceptor command to an angle of attack reference value; wherein if ahigh attitude protection is engaged, said first transformation routineis operable to transform said inceptor command to a pitch anglereference value; and wherein said transformation performed by saidsecond transformation routine depends on what type of protection isengaged and is different for each type of protection.
 2. The flightcontrol apparatus of claim 1, wherein said sensors are operable to senseat least angle of attack, angle-of-attack rate, mach number, flapposition, landing gear position and ice condition of the flight vehicle,said apparatus further including: a calculation routine associated withthe processor to compute an angle-of-attack threshold equal to anangle-of-attack plus a bias based at least in part on an angle-of-attackrate; a computation routine associated with the processor to compute anangle-of-attack reference value based at least in part on a mach number,a flap position, a landing gear position and an ice-condition;comparison routine associated with the processor to compare theangle-of-attack threshold to an angle-of-attack reference value, whereinan automatic function is used by the processor to command the actuatorif the angle-of-attack threshold is larger than the angle-of-attackreference value.
 3. The flight control apparatus as claimed in 1,wherein said sensors are operable to sense at least an airspeed, anairspeed rate and a flap position of the flight vehicle, said apparatusfurther including: a first computation routine associated with saidprocessor to compute an airspeed threshold equal to the airspeed minus abias based at least in part on the airspeed rate; a second computationroutine associated with said processor to compute an airspeed referencevalue based on the flap position of the flight vehicle; and comparisonroutine associated with the processor to compare the airspeed thresholdto the airspeed reference value, wherein an automatic function is usedby the processor to command the actuator if the airspeed threshold islower than the airspeed reference value.
 4. The flight control apparatusof claim 3, wherein said sensors are operable to further sense at leastthe throttle lever position of the flight vehicle, said apparatusfurther comprising: a position sensing routine associated with saidprocessor to process the throttle lever position to define the envelopeof thrust allowable in a predetermined flight control mode; and aprevention routine associated with the processor to prevent an automaticfunction from commanding the actuator if the throttle lever positionindicates that the flight vehicle is not in the allowable thrustenvelope.
 5. The flight control apparatus of claim 1, wherein saidsensors are operable to sense at least a pitch attitude, a pitchattitude rate, a flap position and a flight vehicle height above ground;a first computation routine associated with said processor to compute apitch attitude threshold plus a bias based at least in part on the pitchattitude rate; a second computation routine associated with saidprocessor to compute a pitch attitude reference value based at least inpart on the flap position and the flight vehicle height above ground;and comparison routine associated with the processor to compare thepitch attitude threshold to the pitch attitude reference value, whereinan automatic function is used by the processor to command the actuatorif the pitch attitude threshold is larger than the pitch attitudereference value.
 6. The flight control apparatus of claim 1, whereinsaid sensors are operable to sense at least flap position, landing gearposition, Mach number and altitude; and the feedback, feed-forward andintegral gains are adjusted according to the flap position, landing gearposition, Mach number and altitude of the flight vehicle.
 7. The flightcontrol apparatus of claim 1 comprising: a first gain adjustment routineto adjust an airspeed setting of feed-forward, feedback and integralgains when an airspeed threshold is lower than an airspeed referencevalue; a second gain adjustment routine to adjust an angle-of-attacksetting of feed-forward, feedback and integral gains when anangle-of-attack threshold is greater than an angle-of-attack referencevalue; a third gain adjustment routine to adjust a pitch attitudesetting of feed-forward, feedback and integral gains when a pitchattitude threshold is greater than a pitch attitude reference value; andwherein an automatic function is used by the processor to command theactuator according to the angle-of-attack settings of gains if theangle-of-attack threshold is larger than its reference value, accordingto the airspeed setting of gains if the airspeed threshold is lower thanits reference value, or according to the pitch attitude gains when thepitch attitude threshold is greater than its reference value. 8-23.(canceled)